The Goal-Oriented methodology (GO methodology) is an effective method for the reliability analysis of complex systems. It is especially suitable for the reliability analysis of multi-state complex systems containing the actual logistics, such as current, airflow and liquid flow. In order to solve the limitation that the GO methodology is not suitable for the reliability analysis of the system with feedback loop, the Boolean algebra idea is introduced to construct the Boolean operation formula of the feedback loop. In this paper, a certain type of civil aircraft flap hydraulic system with feedback loop is taken as the research object. According to the structural schematic diagram of the flap hydraulic system, the GO model of the flap hydraulic system is established. Next, the GO calculation is carried out to obtain the reliability of the flap hydraulic system. The comparison between the system reliability without feedback loop and that with feedback loop proves that the GO methodology with feedback loop is more accurate. The reliability of the system is analyzed by using the fault tree analysis (FTA) method, and the GO methodology with feedback loop is compared with the FTA method to verify the availability and correctness of the GO methodology in the reliability analysis and safety evaluation of aircraft flap hydraulic system.

The European Space Operations Centre currently operates five Copernicus Sentinel satellites in the framework of Europe’s Copernicus Earth observation programme. The routine operations rely on a daily orbit determination, carried out on-ground, consisting in a least-squares fit of a dynamical model to GPS navigation solutions generated on-board. The purpose of this paper is the estimation of realistic uncertainties on this daily determined state vector. By comparison with the orbit derived by Precise Orbit Determination, we estimate the 1-sigma errors at approximately 0.5m and 0.5mm/s. Non-stationary errors in the navigation solution preclude their characterisation with a constant covariance matrix. Error whitening is achieved by decreasing the signal-to-noise ratio in the errors through the use of underestimated weights on the data. The approach keeps the errors on the derived state vector unchanged and allows the covariance on the state vector to become realistic.

This article presents a development of a simple analytical aerodynamic model capable of describing the effect of leading-edge vortices (LEVs) on the lift of rotating samara wings. This analytical model is based on the adaptation of Polhamus’ method to develop a sectional two-dimensional lift function, which was implemented in a numerical blade element model (BEM) of a rotating samara blade. Furthermore, wind tunnel experiments were conducted to validate the numerical BEM and to assess the validity of the newly developed analytical lift function. The results showed good agreement between the numerical model and the experimental measurements of rotational speed and rate of descent of the samara wing. The results were also compared with numerical predictions using BEM but adopting different lift coefficient expressions available in literature. This research contributed towards efficient aerodynamic modelling of the lift generated by LEVs on rotating samara wings for performance prediction, which could potentially be used in the design of bio-inspired rotary micro-air vehicles.

Gas turbine blades feature multi-pass internal cooling channels, through which relatively colder air bled from the compressor is routed to cool internal walls. Under rotation, due to the influence of Coriolis force and centrifugal buoyancy, heat transfer at the trailing side enhances and that at the leading side reduces, for a radially outward flow. This non-uniform temperature distribution results in increased thermal stress, which is detrimental to blade life. In this study, a rotation configuration is presented which can negate the Coriolis force effect on heat and fluid flow, thereby maintaining uniform heat transfer on leading and trailing walls. A straight, smooth duct of unit aspect ratio is considered to demonstrate the concept and understand the fluid flow within the channel and its interaction with the walls. The new design is compared against the conventional rotation design. Numerical simulations under steady-state condition were carried out at a Reynolds number of 25000, where the Rotation numbers were varied as 0, 0.1, 0.15, 0.2, 0.25. Realisable version of k-$\varepsilon$ model was used for turbulence modelling. It was observed that new rotation (parallel) configuration’s heat transfer on leading and trailing sides were near similar, and trailing side was marginally higher compared to leading side. An interesting phenomenon of secondary Coriolis effect is reported which accounts for the minor differences in heat transfer augmentation between leading and trailing walls. Due to centrifugal buoyancy, the fluid is pushed towards the radially outward wall, resulting in a counter-rotating vortex pair, which also enhances the heat transfer on leading and trailing walls when compared to stationary case.

One of the most critical aspects of a small satellite is the electrical power system (EPS) since the electrical power is necessary for the satellite to operate correctly during its predefined lifetime. The electrical power system consists mainly of solar cells, batteries, voltage converters and protection circuits. The electrical power system is responsible of providing stable power to the rest of the satellite subsystems.

In satellite electrical power systems, overcurrent protection is now becoming an important function handled by the power distribution module (PDM). This paper proposes a method to evaluate the suggested protection. With the proposed procedure we should be able to verify that every possible failure does not travel through the EPS and cause a fatal degradation of the electrical power system. This will allow a complete evaluation of functionality of the protection hardware.

This paper discusses the design and implementation of the power distribution module (PDM) for the coming generation of small satellites for the Algerian Space Agency (ASAL). The design must provide a reliable protection for the subsystems from the overcurrent associated with a device failure.

Targeting higher efficiencies and lower fuel consumption of turbomachines, heat transfer and profile loss are research topics of particular interest. In contrast to that, the interaction of both was, so far, rarely investigated, but gains in importance in recent research activities. The profile loss of engine components can be characterised by the airfoil wakes at the blade rows utilising established measurement and evaluation methods for which an adiabatic flow is typically supposed. To enable the investigation of the influence of heat transfer at the blade on the loss characteristics, a novel evaluation procedure was set up. In addition to the pneumatic data, the total temperature in the airfoil wake at a linear cascade was measured by means of a five-hole probe with an integrated thermocouple. For the evaluation and analysis of these data, different definitions of the loss coefficient were investigated and, finally, extended to account for thermal aspects. Furthermore, established techniques to average the local wake data were applied and compared with special focus to their suitability for non-adiabatic cases. Moreover, an extended version of the mixed-out average as defined by Amecke was utilised applying not only a far-reaching consideration of a temperature gradient but also the inclusion of the third spatial dimension to enable the evaluation of field traverses in addition to single wake traverses. These techniques were applied to wake measurement data from a linear compressor cascade gained in a special test set-up in the high-speed cascade wind tunnel for different operating points and different blade temperatures. The suitability of the new methods could be proven, and initial steps of the aerodynamic analysis of the resulting data are presented. Thereby, the acquired techniques turned out as powerful methods for the evaluation of wake traverses on compressor and turbine cascades under non-adiabatic conditions.

Aircraft conflict resolution is an important part of air traffic control operations. This study presents a mixed integer linear programming model (MILP) using a space discretisation technique to deal with aircraft conflict resolutions in en-route flight operations. The purpose of space discretisation is to concentrate on only the significant points of the airspace. The model integrates the multi entry point approach with an airspeed adjustment technique in the horizontal plane. The model aims to generate conflict-free trajectories while minimising the total changes in entry points and airspeed values. A new heuristic algorithm was developed due to the complexity of the problem. The computational results demonstrated that the proposed approach resolved aircraft conflicts for 450 different traffic scenarios in less than a minute. Considerable fuel savings were achieved with no significant increase in delay or flight time compared to conventional vectoring techniques in a fixed entry point airspace structure.

Shock control bumps can be used to control and weaken the shock waves that form on engine intakes at high angles of attack. In this paper, it is demonstrated how shock control bumps applied to an engine intake can reduce or eliminate shock-induced separation at high incidence, and also increase the incidence at which critical separation occurs. Three-dimensional Reynolds-average Navier–Stokes (RANS) simulations are used to model the flow through a large civil aircraft engine intake at high incidence. The variation in shock strength and separation with incidence is first studied, along with the flow distribution around the nacelle. An optimisation process is then employed to design shock control bumps that reduce shock strength and separation at a fixed high incidence condition. The bump geometry is allowed to vary in shape, size, streamwise position and circumferential direction around the nacelle. This is shown to be key to the success of the shock control geometry. A further step is then taken, using the optimisation methodology to design bumps that can increase the incidence at which critical separation occurs. It is shown that, by using this approach, the operating range of the engine intake can be increased by at least three degrees.

Distributed real-time avionics networks have been subjected to a great evolution in terms of functionality and complexity. A direct consequence of this evolution is a continual growth of data exchange. AFDX standardised as ARINC 664 is chosen as the backbone network for those distributed real-time avionics networks as it offers high throughput and does not require global clock synchronisation. For certification reasons and engineering research, a deterministic upper bound of the end-to-end transmission delay for packets of each flow should be guaranteed. The Forward Approach (FA) is proposed for the computation of the worst-case end-to-end transmission delay. This approach iteratively estimates the maximum backlog (amount of the pending packets) in each visited switch along the transmission path, and the worst-case end-to-end transmission delay can be computed. Although it is pessimistic (overestimated), the Forward Approach can provide a tighter upper bound of the end-to-end transmission delay while considering the serialisation effect. Recently, our research finds the computation of the serialisation effect might induce an optimistic (underestimated) upper bound. In this paper, we analyse the pessimism in the Forward Approach and the optimism induced by the computation of the serialisation effect, and then we provide a new computation of the serialisation effect. We compare this new computation with the original one, the experiments show that the new computation reduces the optimism and the upper bound of the end-to-end transmission delay can be computed more accurately.

The coaxial compound helicopter with lift-offset rotors has been proposed as a concept for future high-performance rotorcraft. This helicopter usually utilizes a variable-speed rotor system to improve the high-speed performance and the cruise efficiency. A flight dynamics model of this helicopter associated with rotor speed governor/engine model is used in this article to investigate the effect of the rotor speed change and to study the variable rotor speed strategy. Firstly, the power-required results at various rotor rotational speeds are calculated. This comparison indicates that choice of rotor speed can reduce the power consumption, and the rotor speed has to be reduced in high-speed flight due to the compressibility effects at the blade tip region. Furthermore, the rotor speed strategy in trim is obtained by optimizing the power required. It is demonstrated that the variable rotor speed successfully improves the performance across the flight range, but especially in the mid-speed range, where the rotor speed strategy can reduce the overall power consumption by around 15%. To investigate the impact of the rotor speed strategy on the flight dynamics properties, the trim characteristics, the bandwidth and phase delay, and eigenvalues are investigated. It is shown that the reduction of the rotor speed alters the flight dynamics characteristics as it affects the stability, damping, and control power provided by the coaxial rotor. However, the handling qualities requirements are still satisfied with different rotor speed strategies. Finally, a rotor speed strategy associated with the collective pitch is designed for maneuvering flight considering the normal load factor. Inverse simulation is used to investigate this strategy on maneuverability in the Push-up & Pull-over Mission-Task-Element (MTE). It is shown that the helicopter can achieve Level 1 ratings with this rotor speed strategy. In addition, the rotor speed strategy could further reduce the power consumption and pilot workload during the maneuver.

The increasing use of unmanned aerial vehicles in areas such as rescue, mapping, and transportation have made it necessary to study more accurate techniques for calculating flight time estimates. Such calculations require knowing the battery discharge profile. Simplified flight time calculation methods provide data with uncertainties as they are based solely on manufacturer datasheet information. This study presents a setup to measure the battery discharge curve using a LabVIEW interface with a low-cost acquisition system. The acquired data passes through a nonlinear optimisation algorithm to find the battery coefficients, which enables the more precise estimation of its range and endurance. The great advantage of this model is that it makes it possible to predict how the battery will discharge at different rates using just one experimental curve. The methodology was applied to three different batteries and the model was validated with different discharge rates in a controlled environment, which resulted in endurance lower than 3.0% for most conditions and voltage estimation error lower than 3.0% in operational voltage. The work also presented a methodology for estimating cruise time based on the current used during each flight stage.

The prime focus of this work is to estimate stability and control derivatives of an airship in a completely nonlinear environment. A complete six degrees of freedom airship model has its aerodynamic model as nonlinear functions of angle of attack. Estimating the parameters of aerodynamic model in a nonlinear environment is challenging as it demands an exhaustive dataset that could cover the entire regime of operation of airship. In this work, data generation is achieved by simulating the mathematical model of airship for different trim conditions obtained from continuation analysis. The mathematical model is simulated using predicted parameter values obtained using DATCOM methodology. A modular neural network is then trained using back-propagation and Adam optimisation algorithm for each of the aerodynamic coefficients separately. The estimated nonlinear airship parameters are found to be consistent with the DATCOM parameter values which were used for open-loop simulation. This validates the proposed methodology and could be extended to estimate airship parameters from real flight data.

The ability of microramps to control shock - boundary layer interaction at the vicinity of an axisymmetric compression corner was investigated computationally in a Mach 4 flow. A cylinder/flare model with a flare angle of 25° was chosen for this study. Height (h) of the microramp device was 22% of the undisturbed boundary layer thickness (δ) obtained at the compression corner location. A single array of these microramps with an inter-device spacing of 7.5h was placed at three different streamwise locations viz. 5δ, 10δ and 15δ (22.7h, 45.41h and 68.12h in terms of the device height) upstream of the corner and the variations in the flowfield characteristics were observed. These devices modified the separation bubble structure noticeably by producing alternate upwash and downwash regions in the boundary layer. Variations in the separation bubble’s length and height were observed along the spanwise (circumferential) direction due to these devices.

One of today’s most propitious immersive technologies is virtual reality (VR). This term is colloquially associated with headsets that transport users to a bespoke, built-for-purpose immersive 3D virtual environment. It has given rise to the field of immersive analytics—a new field of research that aims to use immersive technologies for enhancing and empowering data analytics. However, in developing such a new set of tools, one has to ask whether the move from standard hardware setup to a fully immersive 3D environment is justified—both in terms of efficiency and development costs. To this end, in this paper, we present AeroVR—an immersive aerospace design environment with the objective of aiding the component aerodynamic design process by interactively visualizing performance and geometry. We decompose the design of such an environment into function structures, identify the primary and secondary tasks, present an implementation of the system, and verify the interface in terms of usability and expressiveness. We deploy AeroVR on a prototypical design study of a compressor blade for an engine.

This paper has proposed a new robust hybrid nonlinear guidance law, which accounts for a missile’s terminal line-of-sight (LOS) angle constraint, in order to intercept a non-cooperative maneuvering target. The proposed hybrid nonlinear guidance strategy consists of two phases; in the first phase, a guidance law named PIGL is derived from prescribed performance control and the inertial delay control method. In PIGL, a revised prescribed performance function is put forward, and a prescribed performance controller with unknown uncertainties is then derived. The controller smoothly drives both the LOS angle and its rate to a predesigned small region under unknown uncertainties that are induced by target’s maneuvers within a fixed time. Then, a guidance law named SIGL is activated, which is derived from sliding mode control and inertial delay control. By driving the desired sliding mode variable to zero within a finite time, the SIGL guidance law is able to achieve high terminal interception accuracy. The robustness of both of the proposed sub-guidance laws has been proved explicitly in this paper. The hybrid guidance law has the advantage of a tunable convergence rate of the LOS angle and the rate of the LOS angle at the beginning period, by which an excessive large initial maneuver can be avoided. Meanwhile, the hybrid guidance law also has the advantage of lower sensitivity to errors in the estimation of the time-to-go.

The aeroelastic phenomenon of limit-cycle oscillations (LCOs) is analysed using a projection-based reduced-order model (PROM) and Navier–Stokes computational fluid dynamics (CFD) in the time domain. The proposed approach employs incompressible Navier–Stokes CFD to construct the full-order model flow field. A proper orthogonal decomposition (POD) of the snapshot matrix is conducted to extract the POD modes and corresponding temporal coefficients. The POD modes are directly projected to the incompressible Navier–Stokes equation to reconstruct the flow field efficiently. The methodology is applied to a plunging cylinder and an aerofoil undergoing LCOs. This scheme decreases the computational time while preserving the capability to predict the flow field accurately. The ROM is capable of reducing the computational time by at least 70% while maintaining the discrepancy within 0.1%. The causes of LCOs are also investigated. The scheme can be used to analyse non-linear aeroelastic phenomena in the time domain with reduced computational time.

Experimental analyses of synthetic jet control (SJC) effects on aerodynamic characteristics of rotor in steady state and in hover were conducted. To ensure the structural strength of rotor and enough interior space for holding the synthetic jet actuators (SJAs), a particular blade with a frame-covering structure was designed and processed, and the experiment was conducted with low free stream velocities and rotor rotation speeds. There were three test conditions. In steady state, there were three free stream velocities (10m/s, 15m/s and 20m/s). In hover state, the rotor was worked with two rotation speeds of 180RPM and 240RPM. In forward flight, the rotor was worked with a rotation speed of 180RPM and a free stream velocity of 7.5m/s. To measure the synthetic jet control effect on rotor in stall, the range of collective pitch was set from 10° to 28° in steady state. The aerodynamic forces and sectional velocity field were measured by using the six-component balance and the Particle Image Velocimetry (PIV) system in the wind tunnel. Flow control effects on the blade based on the synthetic jets (SJ) were experimentally investigated with different jet parameters, such as jet locations, jet angles, and jet velocities. In steady state, the jet closer to the leading edge, and the jet angle of 90° had more advantages in improving the aerodynamic characteristics. Furthermore, the aerodynamic forces and sectional velocity field measurement of rotor in hover were conducted, it showed that SJAs could increase flow velocity at the upper surface, which led to lower upper surface pressure. As a result, the normal forces of rotor with two rotation speeds were increased significantly. These results indicated that the synthetic jet has a capability of increasing the normal force and delaying or preventing the stall of rotor.

In demand of simpler and alternative ground flutter test, a new technique that emulates flutter on the ground has recently emerged. In this paper, an improvement of the test technique is made and verified through the experimental work. The technique utilizes general ground vibration test (GVT) devices. The key idea is to emulate the distributed unsteady aerodynamic force by using a few concentrated actuator forces; referred to as emulated flutter test (EFT) technique. The EFT module contains two main logics; namely, real-time aerodynamic equivalent force calculator and multi-input-multi-output (MIMO) force controller. The module is developed to emulate the subsonic, linear flutter on a specified target structure, which is a thin aluminum clamped-plate with aspect ratio (AR) of 2.25. In this study, doublet hybrid method (DHM) was applied to model the subsonic aerodynamic force, which restricts the application to a 2-dimensional structure. Given that, correlation of several experimental works, such as wind-tunnel flutter test, EFT using laser displacement sensor (LDS), and EFT using accelerometer, on the target structure are investigated to verify the technique. In addition to the flutter boundary, flutter mode shape and trend of aerodynamic damping effect are also presented in this work. Together with these various kinds of test results, application of more compact actuator and an accelerometer as a sensor, makes the current technique the most advanced ground flutter emulation test method.

The research paper addresses the problem of estimating aerodynamic parameters using a Gauss-Newton-based optimisation method. The process of the optimisation method lies on the principle of minimising the residual error between the measured and simulated responses of the system. Usually, the simulated response is obtained by integrating the dynamic equations of the system, which is found to be susceptible to the initial values, and the integration method. With the advent of the feedforward neural network, the data-driven regression methods have been widely used for identification of the system. Among them, a variant of feedforward neural network, extreme learning machine, which has proven the performance in terms of computational cost, generalisation, and so forth, has been addressed to predict the responses in the present study. The real flight data of longitudinal and lateral-directional motion have been considered to estimate their respective aerodynamic parameters. Furthermore, the estimates have been validated with the values of the classical estimation methods, such as the equation-error and filter-error methods. The sample standard deviations of the estimates demonstrate the effectiveness of the proposed method. Lastly, the proof-of-match exercise has been conducted with the other set of flight data to validate the estimated parameters.

To establish flutter onset boundaries on the flight envelope, it is required to determine the flutter onset dynamic pressure. Proper selection of a flight flutter prediction technique is vital to flutter onset speed prediction. Several methods are available in literature, starting with those based on velocity damping, envelope functions, flutter margin, discrete-time Autoregressive Moving Average (ARMA) modelling, flutterometer and the Houbolt–Rainey algorithm. Each approach has its capabilities and limitations. To choose a robust and efficient flutter prediction technique from among the velocity damping, envelope function, Houbolt–Rainey, flutter margin and auto-regressive techniques, an example problem is chosen for their evaluation. Hence, in this paper, a three-degree-of-freedom model representing the aerodynamics, stiffness and inertia of a typical wing section is used(1). The aerodynamic, stiffness and inertia properties in the example problem are kept the same when each of the above techniques is used to predict the flutter speed of this aeroelastic system. This three-degree-of-freedom model is used to generate data at speeds before initiation of flutter, during flutter and after occurrence of flutter. Using these data, the above-mentioned flutter prediction methods are evaluated and the results are presented.